Cooling of hot surfaces with cold fluid film is commonplace in many engineering problems including vertical takeoff and landing (V/STOL) and gas turbine blades. For example, gas turbines require proper cooling mechanism to protect its blades from thermal stresses due to hot combustion gases. The problem becomes aggravated by the need for higher turbine inlet temperature to generate more power. Film cooling is used as a mechanism for reducing such thermal stress and for increasing lifetime for a turbine blade. It commonly works in the form of a row of holes located in the spanwise direction, through which cold jets are issued into the hot crossflow. The penetration of cold jet into the main flow creates a three-dimensional flow field. Thus, the trajectory and physical path of thermal jet and the mixing mechanism of jet in the crossflow are critical design parameters.
FIG. 1 A shows a general schematic of the tip of a turbine blade with film cooling holes and coolant plenum. A simplified form of this curved surface may be considered as a flat plate with a single round jet injected in the crossflow at an angle α. This geometry has been extensively investigated for cooling performance for a wide range of blowing ratio (i.e., momentum ratio of injected air to crossflow). FIG. 1B shows a schematic control volume of hot air passing over a flat surface (e.g., a turbine blade). This surface of study has a row of injection holes through which the cool air is issued at an angle α=35°. The cool jet at temperature Tj=150K is injected into the hot freestream of Tfs=300K. The injection ducts are circular pipes with diameter equal to d=2.54 mm. The injection hole fonned by the intersection of the injection pipe with the wind tunnel is an ellipse with the minor and the major axes d and D=d/(sin α), respectively. The distance between the hole centers is L=3d. The selected mean flow velocities, static pressures and temperatures (i.e., densities) in the injection pipe and the wind tunnel gives a blowing ratio M=1. The inlet section is located at x=−20d and the exit at x=29d. The other dimensions and boundary conditions are shown in FIG. 1B. The flat (blade) surface is considered adiabatic.
At the freestream inlet x=−20d, an injected mass flow rate inlet condition was applied with the density ratio of ρj/ρf=2, velocity ratio of uj/uf=0.5, and turbulent intensity of 5%. At the exit plane x=29d, the gauge pressure at the outlet boundary is maintained at 0 Pa. The work surface is an adiabatic wall with a single row of holes through which cool air at temperature is equal to Tj=150K is injected at an angle of a=35° into the hot freestream of temperature Tf=300K. The domain extends from the plenum base at y=−6d to y=20d from work surface where a pressure-far-field boundary condition was applied. The periodic boundary condition was applied in the crosswise direction (at z=±1.5d) in the computational domain. Periodic boundary conditions were also employed in the spanwise direction, on all sides of the plenum.
Despite various innovative techniques, the film cooling effectiveness is ultimately limited by the loss of flow attachment just downstream of the hole. This is due to the “lift-off” of the cold jet beyond a threshold momentum ratio, as shown in FIGS. 2A-2C. The film cooling configuration which is designed for peak flow performance may not actively regulate itself for off-design conditions.
In modern turbomachinery applications (e.g., next generation aircrafts), the hot gas path temperatures inside the turbine are significantly impacted by efforts to reduce noise, fuel burn and emissions. This necessitates significant advancements in blade cooling technologies. It is known that cool jets parallel to the blade surface are the best possible design for film cooling. However, this has generally been infeasible from a manufacturing standpoint and industry resorts to the next best option, namely the shaped holes.
In the near field of the film cooling jet, the dynamic large scale structures control the mixing process. This three-dimensional mixing shown in FIGS. 2A-2C, from detached eddy simulation (DES) of film cooling, determines the normal and transverse penetration of the jet. In FIG. 2A, counter rotating vortical structures at the jet exit plane are colored by the static temperature. In addition, in FIG. 2D, a schematic of a jet-crossflow interaction and a Laser-induced fluorescence (LIF) image description are given. Accurate actuation of such flow field will greatly influence the near wall heat transfer process and the film cooling effectiveness. The complex dynamic nature of the spanwise vortices makes it desirable to use an active mode of control that will interact with the flow field temporally and spatially in the near wall region.
Accordingly, there is a need in the art for a method and apparatus for more efficiently cooling turbine blade.